1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a turbine airfoil with a showerhead leading edge cooling arrangement.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine includes a turbine with one or more stages of rotor blades and stator vanes used to convert energy from the hot gas flow into mechanical work to drive the compressor or, in the case of an industrial gas turbine engine (IGT) to drive an electric generator. In order to extract the highest amount of mechanical energy, the hat gas flow is passed into the turbine having the highest possible temperature. However, the highest possible temperature at the turbine inlet is limited to the material capabilities of the first stage vanes and blades. if the blades or vanes become too hot, the parts can fail.
In order to allow for higher gas flow temperatures, the turbine airfoils are cooled by passing a pressurized cooling air through the internal passages formed within the airfoils. A combination of internal convection cooling with impingement and film cooling is sued in order to maximize the cooling ability while minimizing the amount of pressurized cooling air used.
The leading edge of the stator vanes and the rotor blades are exposed to the highest gas flow temperature and therefore require the highest amount of cooling to prevent hot spots. In the prior art, an airfoil leading edge is cooled with backside impingement in series with showerhead film cooling. showerhead film rows are supplied cooling air from a common impingement cavity and discharge the cooling air at various gas side pressures. U.S. Pat. No. 6,099,251 issued to LaFleur on Aug. 8, 2000 and entitled COOLABLE AIRFOIL FOR A GAS TURBINE ENGINE shows this type of leading edge cooling arrangement. As a result of this method for cooling the leading edge, cooling flow distribution and pressure ratio across the showerhead film holes for the pressure side and suction side film row is predetermined by the impingement cavity pressure. Also, the standard film slots pass straight through the airfoil wall at a constant diameter and exit at an angle to the surface. Some of the coolant is subsequently injected directly into the mainstream causing turbulence, coolant dilution and a loss of downstream film cooling effectiveness. And, the film slot breakout on the airfoil surface may induce stress problem in a blade cooling application.
U.S. Pat. No. 3,819,295 issued to Hauser et al on Jun. 25, 1974 and entitled COOLING SLOT FOR AIRFOIL BLADE discloses a turbine blade with a trailing edge cooling passage formed by rows of holes drilled at about 90 degrees offset to form square shaped nodes that act as turbulence promoters to the cooling air flow. One problem with the Hauser et al invention is that the passages are formed by drilling holes in the trailing edge blade material. Because the passages are formed by drilling holes, the passages for the cooling air do not flow in a serpentine path as is provided for in the present invention. Also, the drilled holes cannot be formed close to an end of the blade. The holes on the tip of the blade in Hauser et al have to be drilled from the tip and not from the trailing edge. On both the tip and the bottom of the trailing edge in the Hauser et al invention, the holes for the cooling air to pass are not formed near the inner or outer extremes because the holes cannot be drilled without passing through the material on the extremes as is accomplished in the present invention and described below. Thus, the method of forming cooling air passages cannot be used to form cooling holes for the leading edge showerhead of the turbine airfoil as is the case for the present invention.